1. Field of the Invention
This invention relates to means for directing cooling air to critical parts of hot section turbine blades in gas turbine engines.
2. Description of the Prior Art
In the course of gas turbine engine development, tremendous effort has been directed at raising the internal operating temperatures of such engines to improve thermodynamic efficiency. As turbine inlet temperatures have been increased in pursuit of this goal, it has become necessary to provide cooling air to hot section turbine blades and vanes in order to limit temperatures of those components to levels that can be accommodated by the blade and vane materials. The air that is used for this cooling function is usually compressed to pressures that meet or exceed the gas pressures inside the turbine section. Because the air has undergone the work necessary for compression, this cooling air must be used as efficiently as possible to limit the power required by the engine's compressor section in order to compress that air. To limit the amount of cooling air used, intricate cooling air flowpaths and passages are utilized that are intended to use the cooling air in a highly efficient manner.
In smaller airflow size engines, blade cooling configurations are generally restricted to fairly simple designs because of small dimensions and limitations of current manufacturing technologies. The implication is a typical smaller engine turbine blade or vane cannot be provided with the highly complex, internal air cooling passage configuration typically used today in larger gas turbine engines.
One particular problem with smaller engines is that tip sections of turbine blades are extremely difficult to cool efficiently. The cooling air used to internally cool turbine blade tips has increased its temperature by thermal pickup in the lower portion of the blade rendering it less effective for cooling purposes. In downstream sections of the turbine blade tips some of the cooling air has been bled out of the trailing edge cooling holes before it reaches the blade tip region, thereby reducing cooling air velocity and, consequently, its cooling effectiveness. Adding to these difficulties of cooling small turbine blades, the downstream trailing edge of the blade tip region is usually very thin for aerodynamic performance reasons, which limits the ability to duct cooling air into this region.
As a result of these inherent limitations, design cycle temperatures of these small engines are restricted and engine performance is thereby limited. Further, the turbine blade tips often become a life-limiting engine component problem area. As the turbine tips deteriorate, due to oxidation and corrosion accumulating during engine use, the engine performance drops below minimum acceptable levels. The engine must then be removed from the aircraft and the turbine section refurbished. Maintenance and overhaul of the turbine section to correct deteriorated blade tips is both expensive and time consuming.
It is, therefore, an object of the present invention to provide a means for cooling tips of turbine blades in turbine sections of gas turbine engines with a system that can be utilized in relatively small engine configurations.
Another object of the present invention is to provide a source of cooling air that can be directed specifically to turbine blade tips in a turbine section of a small gas turbine engine.
It is another object of the present invention to provide a film of coating air along a radially outer most wall of a turbine section of a small gas turbine engine for the purpose of cooling turbine blade tips with a limited amount of cooling air.
These and other objects will become more readily apparent upon reference to the following description in conjunction with the appended drawings.